The present invention relates to a tip turbine engine, and more particularly to a hollow fan blade with a diffuser that turns the flow within a constrained radial distance while maintaining a balanced mass.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor and ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine which rotatably drives the high pressure compressor through the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine which rotatably drives the fan and low pressure compressor through a low pressure shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in an elongated engine structure of considerable longitudinal length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor and ignited to form a high energy gas stream which drives the turbine integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490.
The tip turbine engine provides a thrust to weight ratio equivalent to conventional turbofan engines of the same class within a package of significantly shorter length.
The tip turbine engine utilizes hollow fan blades as a centrifugal impeller. Axial airflow from an upstream source such as ambient or an axial compressor must be turned into a radial airflow for introduction into the hollow fan blades. Communicating airflow through the hollow fan blades within a relatively limited radial turning distance of a fan turbine rotor without excessive losses provides an engine design challenge.
Hollow bypass fan blades such as those generally disclosed in U.S. Patent Application Publication Nos.: 20030192303, 20030192304, and 20040025490 may form a mass imbalance about a fan blade outer periphery. During rotational loading, the mass imbalance about the fan blade outer periphery may generate a moment on the fan blade leading to excessive bending stresses.
Accordingly, it is desirable to provide a lightweight hollow fan blade with a core airflow passage which turns an airflow within a minimal radial turning distance without excessive losses while assuring a balanced tip shroud mass.